Enabling a spacecraft to embark on a deep-space mission is generally quite a challenge for aerospace designers and engineers, for they must find a way to equip the spacecraft with enough propulsion capability to successfully travel long distances through space and thereby carry out the mission. In taking on the challenge of providing sufficient propulsion, designers and engineers must generally anticipate the overall mass payload likely to be onboard the spacecraft during the mission and the amount of propellant necessary to support such a payload during flight. Depending on the type of mission, the onboard payload itself may include, for example, astronauts, human life support equipment, mission-related tools and hardware, et cetera. During space flight, since the mass of dwindling propellant must also be considered as part of the spacecraft's onboard payload, designers and engineers must generally find a way to equip the spacecraft with the propulsion capability of supporting as much payload as possible with as little propellant as is necessary.
When a conventional chemical propulsion system is proposed for a given spacecraft, a large percentage of the payload mass-carrying capacity onboard the spacecraft is designated for accommodating the propellant. In designating such a large portion of the payload capacity for the propellant, the amount of payload capacity remaining for other mission-critical items is thereby generally reduced. As a result, a proposed space mission may ultimately be deemed infeasible due to payload capacity and cost design constraints. In attempting to address this problem, some studies have shown that increasing the exhaust velocity of a spacecraft's thruster(s) can significantly reduce the amount of propellant required onboard for a given space mission. To date, however, cryogenic chemical propulsion systems incorporated in rockets, for example, have only been able to produce exhaust velocities approaching 5 kilometers per second (km/s), and storable chemical propulsion systems in use onboard other spacecraft have only been able to produce exhaust velocities that are lower still. In light of such, a propulsion system that does not largely rely on energy produced through chemical reactions is instead being sought for utilization onboard a spacecraft intended for deep-space missions.
In contrast to such chemical propulsion systems, electric propulsion systems incorporated within thrusters onboard spacecraft have been shown to produce exhaust velocities on the order of about 10 km/s, or even higher. Thus, in utilizing electric propulsion systems to produce such improved velocities, the amount of propellant necessary for successful deep-space travel is thereby generally reduced. As a result, a smaller percentage of a spacecraft's payload mass-carrying capacity is taken up by the propellant, thereby allowing a larger percentage of the payload capacity to be dedicated to other items necessary for a successful space mission. In light of such, therefore, electric propulsion technology, as opposed to chemical propulsion technology, seems to be a more promising and viable candidate for being incorporated within the propulsion systems of spacecraft intended for carrying out deep-space missions.
In brief, electric propulsion systems generally fall into three main categories. These categories include electrothermal propulsion systems, electromagnetic propulsion systems, and electrostatic propulsion systems. In electrothermal propulsion systems, a propellant undergoes thermodynamic expansion via controlled thermal heating. In this way, the resultant propellant gas is accelerated until it ultimately reaches a certain exhaust velocity as naturally dictated by gas thermodynamics. In electromagnetic propulsion systems, a propellant is initially converted into plasma (i.e., an ionized gas) within, for example, a plasma production chamber. Thereafter, the plasma is accelerated via an electromagnetic field into a high-velocity exhaust stream. In electrostatic propulsion systems, a propellant is initially converted into electrically charged ions (i.e., a plasma) within, for example, an ionization chamber. Thereafter, the charged ions are accelerated via an electrostatic field into a high-velocity exhaust stream.
In recent years, the utilization of electrospray techniques as means for ionizing a liquid propellant and producing charged particles for electric propulsion has received considerable attention. In a conventional electrospray technique, a slightly conductive electrolytic liquid is channeled through a capillary needle and emitted from a tip opening in the needle. At the same time, a strong electrostatic field is applied at the tip opening of the needle, thereby causing an imbalance of surface force due to the accumulation of charges on the surface of the emitted liquid. If both the flow rate of the liquid and the electric field at the needle tip opening are maintained at proper levels or strengths, a liquid cone commonly referred to as a “Taylor cone” is thereby formed at the needle tip along with a jet issuing forth from the cone's apex. As the jet travels further away from the Taylor cone, the jet eventually becomes unstable and separates into a spray of charged droplets. In this form, the spray of charged droplets, or “electrospray,” is said to be in a “cone-jet mode.”
In attempting to utilize such electrospray technology for the production of charged particles (i.e., for ionization), some of the inherent benefits generally anticipated and sought after are as follows. First, electrospray ionization can be carried out by utilizing a substantially inert fluid as a propellant. Second, electrospray ionization consumes less energy than more conventional methods of electric propulsion. Third, electrosprays having various charge-to-mass (q/m) ratios can be produced by simply adjusting the flow rate of the liquid propellant and/or the strength of the applied electric field.
To date, some scientific investigations and engineering applications have already demonstrated that electrospray ionization is a suitable means, in certain instances, for producing charged particles for space propulsion. For example, electrospray technology has been utilized in thrusters incorporating electrostatic colloid propulsion systems. In general, a colloid thruster is a specific type of electrostatic thruster that operates by utilizing an electrostatic field to accelerate numerous charged liquid drops (i.e., a colloid beam) emitted from a Taylor cone to thereby generate thrust. In practice, instead of using a single capillary needle, which alone is incapable of producing the required quantity of charged drops or particles necessary for adequate propulsion, an array of emitters consisting of several hundreds of needles is commonly utilized in an individual colloid thruster. When equipped with such emitter arrays, research has shown that colloid thrusters are individually able to deliver thrust levels ranging as high as up to several hundreds of micro-newtons (μN). At such thrust levels, the high-performance propulsion of small spacecraft, including limited translation of small spacecraft through space, is thereby made possible.
In general, electrostatic colloid thrusters incorporating electrospray technology offer many attractive benefits over other electric propulsion system technologies. Some of these benefits include lower energy consumption and higher energy efficiency, which are direct results of alternatively utilizing electrospray technology to ionize propellant. Another benefit is the ability to utilize an inert propellant at ambient temperature levels. As a result of this particular benefit, a less complex and smaller sized propellant storage-and-delivery system may be utilized onboard a spacecraft, thereby improving overall system reliability and also freeing up payload space. Furthermore, still another benefit is flexibility, for colloid thrusters incorporating electrospray technology are able to provide varying thrust levels as well as a broad range of specific impulse (i.e., thrust per unit mass flow of propellant) levels.
Despite such benefits, applications of electrostatic colloid thrusters incorporating electrospray technology have primarily been limited to micro and nano-spacecraft and maintaining the precise positions of such spacecraft in space. Such relegation is mostly due to the low thrust levels and relatively low specific impulse (ISP) levels that have been characteristic of such colloid thrusters heretofore. In particular, studies to date on the feasibility of utilizing electrospray technology onboard spacecraft for propulsion have primarily focused on the hardware (capillary needles) necessary for supporting numerous electrospray-producing liquid Taylor cones on a spacecraft. In brief, such studies have demonstrated that in order to support a large enough number of Taylor cones onboard a spacecraft to sufficiently improve thrust, the number density of capillary needles onboard a spacecraft (i.e., the number of needles per unit area) must be increased so that thousands of needles can be integrated into the spacecraft's thruster(s). Because of inherent onboard space and payload limitations, however, the actual number of capillary needles that can be successfully included aboard such a spacecraft is generally somewhat limited, and hence the number of electrospray-producing Taylor cones that can be sustained onboard is correspondingly limited as well. As a result, both the propellant mass flow rate and the level of thrust that can be achieved by such a spacecraft are also limited. Furthermore, such studies have also demonstrated that a liquid Taylor cone, when operating in the cone-jet mode by means of a capillary needle channeling electrolytic fluid in the presence of an electrostatic field, tends to produce a liquid jet that is too stable and thus generally unable to quickly separate into a spray of charged droplets. In an attempt to counteract and reduce such stability, a very high onset voltage (VON) is often applied between the capillary needle and an electrically conductive extractor so as to successfully extract colloid beams of discrete droplets from the Taylor cone at the needle tip's opening and thereby produce an electrospray. When such a large onset voltage is applied as such, however, the liquid Taylor cone frequently emits solvated ions (i.e., ions with water molecules attached thereto) along with the charged droplets. In general, these solvated ions characteristically have much higher charge-to-mass ratios (q/m) than the charged droplets. As a result, an overall electrospray of particles with highly disparate and non-uniform charge-to-mass ratios is ultimately produced, which is generally undesirable in electric propulsion systems. In sum, therefore, the low number density of capillary needles that can be integrated into a spacecraft's thruster(s) and also the characteristic stability of liquid jets that issue forth from liquid Taylor cones are two primary factors that have undesirably limited the thrusting capabilities, and therefore applications, of colloid thrusters incorporating electrospray technology in recent modern times.
In light of the above, there is a present need in the art for both a method and a thruster that are based on a modified electrospray technology capable of producing large quantities of uniformly charged particles for the high-performance propulsion of spacecraft in and through space.